The present invention relates generally to gas turbine engines and, more specifically, to the cooling of turbine blades or buckets supported on one or more gas turbine rotor wheels.
In a gas turbine engine, air is pressurized in a compressor, mixed with fuel in one or more combustors and ignited to thereby generate hot combustion gases. Energy is extracted from the combustion gases in one or more turbine stages disposed downstream of the combustors.
Each turbine stage includes a stationary turbine nozzle having a row of vanes or blades which direct the combustion gases to a cooperating row of turbine buckets mounted on a wheel fixed to the turbine rotor. The turbine buckets are typically hollow (or cast with internal passages or channels)and are provided with air bled from the compressor (compressor discharge or extraction air) for cooling the buckets during operation.
Bucket airfoils have a generally concave pressure side and an opposite and generally convex suction side extending generally axially between leading and trailing edges, and radially from a platform to an outer tip.
In view of the three-dimensional, complex combustion gas flow distribution over the bucket airfoils, the different portions thereof are subjected to different heat loads during operation. The very high temperatures generate thermal stresses in the airfoils which must be suitably limited in order to prolong the service life of the airfoils and hence the buckets.
The airfoils are typically manufactured from superalloy cobalt- or nickel-based materials having sustained strength under high temperature operation. As noted above, the useful life of the buckets is limited, however, by the maximum stresses and high temperatures experienced by the airfoil portions of the buckets.
Accordingly, the prior art describes various internal cooling channels or circuits, some of which incorporate different forms of heat transfer-increasing turbulator ribs, pins or the like for cooling the various portions of the airfoil.
For example, U.S. Pat. No. 6,174,134 (Lee et al.), assigned to applicant, discloses an airfoil cooling configuration for effecting enhanced cooling of the trailing edge area of the airfoil. Cooling air flowing radially outwardly in a passage adjacent the trailing edge is channeled by multiple crossover holes into a cavity extending along the trailing edge.
In U.S. Pat. No. 6,607,356, the turbine airfoil includes pressure and suction sidewalls having first and second cooling circuits disposed therebetween, separated by a longitudinally, i.e., radially, extending bridge. The aft or trailing edge circuit includes a bridge formed with a row of inlet holes extending along the length of the bridge, allowing radially outwardly-directed flow in one channel of the circuit to crossover into a second channel closer to the trailing edge.
In a continuing search for improved cooling circuits that provide enhanced cooling with efficient use of compressor air, it has been determined an internal bucket cooling circuit that supplies lower temperature cooling medium to the bucket trailing edge region or cavity, and especially to a known hotspot at the outer tip of the trailing edge region would be desirable.